Hybrid rocket motors generally employ two or more initially separate propellants which usually exist in two different phases or physical states; namely, as a fluid propellant and a solid propellant. Preferably, fluid (liquid or gas) oxidizers are used with solid fuel grains, although, solid oxidizers have been used with fluid fuels in what are often referred to as "inverse hybrid rocket motors".
A barrier or valve is used to initially maintain the hybrid propellants in separate storage areas within the hybrid rocket. The fluid propellant is usually contained in a storage tank forward of the main combustion chamber in which the solid propellant grain is fixed. The fluid propellant tank is connected to the main combustion chamber by any of several known methods which, as will be discussed below, often involve valved tubes or pipes.
At start-up, the barrier (whether involving a valve or other means) separating the fluid propellant from the solid propellant is removed and thus, the fluid propellant is allowed to flow freely and usually with high velocity into the main combustion chamber where it combustively reacts with the solid fuel. This combustion generates high temperature, high pressure gases which exit the rocket nozzle and thereby create thrust. Note, depending on the types of oxidizers and fuels used, supplemental igniters and/or "preheaters" may also be necessary to initiate combustion. In other words, not all effective oxidizer-fuel combinations are hypergolically active (spontaneously ignitable upon contact).
Many different methods have been used in previous hybrid rockets to maintain an initial separation between the two propellants. These include, for example: Sessums (U.S. Pat. No. 3,142,152) who teaches a hybrid rocket motor having a non-pyrotechnically activated fluid flow valve located in an isolated compartment between the fluid and solid propellant compartments and further having a subsequent barrier provided by one or more spiraled melt away hoses embedded in the solid grain.
Strauss et al. (U.S. Pat. No. 3,177,657) teach the use of a pair of annular, burstable sealing bands as hybrid propellant barriers. A solid propellant igniter, when fired, causes a high pressure differential that bursts the first sealing band which initially separates the igniter from the liquid propellant. This igniter continues to create gases which pressurize the liquid propellant. When adequately pressurized, the liquid bursts a second annular band that initially separates the liquid from the solid propellant. The liquid is then forced into the combustion chamber where it reacts with the solid propellant to cause thrust.
Yet another pressure sensitive rupturable hybrid barrier is disclosed by Novotny (U.S. Pat. No. 3,325,998). Specifically, Novotny uses pressure sensitive rupture discs to seal the liquid propellant away from the solid propellant until it has been pressurized sufficiently to burst upon start-up.
The general art of rocketry has produced several pyrotechnically activated valve means such as that disclosed by Ferris (U.S. Pat. No. 3,431,731) which has an explosively actuated poppet valve means for rocket motors that involves a frangible diaphragm. Upon firing the poppet, the diaphragm is ruptured and a spring forces the valve closed.
Another explosively activated poppet-type valve means is taught by Hosek (U.S. Pat. No. 3,478,760) and involves the alternate opening or closing of the valve by the firing of opposing "squibs." Squibs are known in the art as comprising combustible materials having rapid burning characteristics which generate high pressure and/or high temperature gas.
The majority of prior art references involving fluid-solid separation means entail valves or other mechanisms used to control flow rate of fluid propellant to the combustion chamber thereby "throttling" the thrust. Examples of these include: Muzzy (U.S. Pat. No. 3,557,556) (Muzzy I) who teaches a hybrid rocket thrust modulator which is operated by a multiplicity of valves. During a low thrust cycle, a single valve feeds a liquid propellant to a single port in the center of the solid propellant grain. Other, radial ports in the solid propellant grain are fed by other valves to increase combustion during higher thrust cycles.
Another hybrid thrust control system is taught by Vickland (U.S. Pat. No. 3,677,011) who utilizes a plurality of solenoid valves in combination with spring forced poppet valves to control at least two fluid propellant flows.
Bennett (U.S. Pat. No. 3,714,783) discloses an inverse hybrid rocket which employs a non-pyrotechnic separation or barrier means involving an electrically controlled rack and pinion assembly.
What in some respects appears to be a combination of many of the above separation or barrier concepts is the hybrid rocket disclosed by Massie (U.S. Pat. No. 3,715,888) which uses a temperature sensitive disintegratable tube exhibiting a plurality of diameters to provide "throttling". This tube is embedded in the solid fuel grain to maintain separation of the fluid from the solid in a manner similar to that of Sessums as described above. Massie also discloses pressure sensitive rupture discs which retain the liquid in separation before start-up. Then, at start-up, the liquid is pressurized to rupture the discs and begin flow to the combustion chamber.
Stone (U.S. Pat. No. 3,789,610) also combines some of the above concepts in teaching an explosively actuated liquid oxidizer poppet valve which, when explosively fired to open, causes an upstream burst diaphragm to rupture and permit oxidizer flow to the combustion chamber. The hybrid liquid flow is terminated by firing an opposing explosive to close the oxidizer poppet.
Holzman (U.S. Pat. No. 4,424,679) involves yet another valve device for separating the fluid propellant from the solid propellant in a hybrid rocket. Holzman includes either the use of a bypass flow in combination with a solenoid valve, or a solenoid controlled valve which uses a spring forced poppet. A feedback transducer is used to control the desired orifice size of the solenoid valve.
Further, the prior art is replete with efforts to improve rocket motor ignition. Dual functions of igniters which must also act as pressure sensitive barrier means have been described. For example, MacDonald (U.S. Pat. No. 2,627,160) teaches an explosive charge igniter that is fixed in a solid propellant rocket nozzle and is employed to both rupture a nozzle closure element and also ignite the solid rocket propellant grain.
Similarly, Heady (U.S. Pat. No. 3,570,405) teaches an igniter charge bonded to an exit nozzle burst disc. Thus, when fired, this device not only ignites a solid propellant grain, it initiates the production of gases which elevates the interior pressure to eventually burst the disc and thereby allow propellant gas exit flow through the exit nozzle.
Burnside (U.S. Pat. No. 3,017,748) teaches the use of a black powder (or other pyrotechnic material) charge to ignite a hybrid rocket motor. At the head end of the combustion chamber, the igniter material is released upon ignition into the combustion chamber to ignite the solid propellant grain. A solid propellant layer, which seals perforations in a fluid propellant tank fixed in the combustion chamber, is then burned away to release the fluid propellant into the combustion chamber to then combust with the solid propellant.
Finally, certain hybrid propellant combinations require preheating of the oxidizer fluid before it enters the combustion chamber. Fluid fuels, such as propane have been burned in the presence of the oxidizer stream in, for example, Holzman (U.S. Pat. No. 4,424,679) and Knuth et al. (U.S. Pat. No. 5,010,730). Solid fuel preheaters have also been described, such as in Bradford (U.S. Pat. No. 3,518,828) who discloses a hybrid ignition system comprising a hollow cylinder of solid propellant contained in a structure apart from the main combustion chamber. This solid propellant cylinder produces fuel-rich, hot exhaust gases which flow into the ignition section of the main combustion chamber and there come into contact with oxidizer vapors flowing from the liquid oxidizer supply. These gases and vapors react to cause ignition of the solid rocket fuel in combination with the incoming liquid oxidizer.
Similarly, Muzzy (U.S. Pat. No. 3,782,112) (Muzzy II) discloses a hybrid rocket injector which gasifies and aerates a liquid oxidizer flowing to the combustion chamber. The injector has a hollow cylindrical gas generating portion of solid propellant between and separate from the main liquid and solid propellant storage chambers. This hollow portion of propellant is maintained in an annular chamber surrounding a tubular liquid oxidizer passageway running to the combustion chamber. This passageway has several openings therein which communicate with the annular chamber. In operation, the portion of propellant in the annular chamber is ignited and gives off hot gases which travel through the openings in the liquid oxidizer passageway and thereby vaporize and aerate the liquid oxidizer as it flows to the combustion chamber.
However, in spite of the many attempts as shown in the prior art, there still remains a need for a simple device which provides a structural barrier between the liquid and solid components of a hybrid rocket motor which is easily operable to remove the barrier from the fluid propellant flow path. A further benefit is realized from a hybrid propellant barrier which also functions as the means for ignition of the hybrid combustion reaction. Moreover, it would be preferable if the device housing the barrier would also provide a fluid oxidizer preheater to increase ignition and combustion efficiency. It is toward the resolution of these previously unfilled needs that the present invention is directed.